It is commonly known that the aeronautical industry requires structures which on one hand can support the loads to which they are subjected, complying with high strength and rigidity requirements, and on the other hand are as light as possible. A result of this requirement is the increasingly extended use of composite materials in primary aircraft structures, involving an important weight saving compared to the use of metallic materials.
The main structure for supporting surfaces of airplanes is formed by a leading edge, a torsion box and a trailing edge. The torsion box of an aircraft is in turn formed by several structural elements. Typically, the process for manufacturing a torsion box is considerably manual and is carried out in a number of steps. The structural elements forming the box are manufactured separately and are mechanically joined with the aid of complicated jigs to achieve the necessary tolerances, which are given by the aerodynamic and structural requirements. This involves different assembly stations and a large amount of joining elements, which entails weight penalties, high production and assembly costs, greater necessary logistic capacity and worse aerodynamic quality in outer surfaces. If the parts are made of composite material, they are manufactured by stacking the different fiber layers and thus forming the desired element layer by layer. At this point, the composite material requires a rather expensive curing process to achieve all its properties.
For this reason, there have recently been great efforts to achieve an increasingly higher level of integration in the production of torsion boxes in composite material and thus prevent the aforementioned drawbacks. The problem consists mainly of generating sufficient pressure in all the elements during the joint curing process.
Thus, there are several known documents describing manufacturing methods which achieve the integration of typical structural elements with the aid of special curing jigs, assembling the remaining elements in the following assembly stages. This is the case of U.S. Pat. No. 5,216,799 (integration of ribs with spars), patent document EP 1074466A1 (integration of ribs) and U.S. Pat. No. 5,735,486 (integration of stringers-skins). Other levels of integration are achieved with the solutions set forth in U.S. Pat. No. 6,237,873B1, describing the manufacture of closed cross-sections and their subsequent joining, and U.S. Pat. No. 6,190,484B1, where contiguous boxes are adjoined to be jointly cured.
Patent documents EP 0582160A1, U.S. Pat. No. 6,896,841B2, U.S. Pat. No. 5,454,895, WO 2004/000643A2 and U.S. Pat. No. 5,817,269 are focused on the jig system for enabling the manufacture of the complete part in a single curing with good quality, whether with jigs which are inflated during the curing or use the difference of thermal expansion of different materials for exerting pressure during the curing at high temperatures.
However, all these solutions start from the basis of individually “pre-stacking” the basic structural elements and curing them jointly with the suitable jigs, which is not a real complete integration, while at the same time the manufacturing costs are high due to the high number of parts to be stacked, there further being a non-uniform passage of loads between the stacked elements.
The present invention is aimed at solving these drawbacks.